1. Field of the Invention
The invention relates to jet propelled aircraft, such as certain types of military aircraft, in which the engine is incorporated directly into the fuselage of the aircraft. In this type of aircraft the rear nozzle of the turbojet engine is situated as a direct extension of the fuselage of the aircraft, and it is necessary to ensure a seal between the nozzle and the rear end of the fuselage.
By way of illustration, FIG. 1 of the drawings shows the rear end of a fighter aircraft 2 in which a turbojet engine 5 for propelling the aircraft is mounted in a housing 4 within the fuselage of the aircraft. The turbojet engine 5 is attached to the housing 4 at the front by pivot means 3 which allows the engine to pivot about a horizontal axis. During operation of the turbojet engine 5, the walls of the engine, as well as the equipment attached to its periphery, can reach temperatures between 250.degree. and 350.degree. C. in the case of military aircraft. It is therefore necessary to protect the aircraft structure at the front of the aircraft from these increases in temperature. To this end, cold air is drawn from the air intake of the aircraft and is led to the inlet of the housing 4 for removing the heat given off during running of the turbojet engine 5. The air current thus created flows inside the housing 4 through a tubular duct 7 formed between the housing 4 and the turbojet engine 5, and is discharged to the exterior of the aircraft at the position of the nozzle constituting the rear end of the aircraft.
This cooling of the engine compartment is therefore carried out by a flow which is the equivalent of two per cent of the flow through the engine itself. Although this cooling flow is low, incorrect discharge thereof can greatly disturb the operation of the thrust nozzle and the aerodynamic behavior of the rear of the aircraft. Accordingly, the cooling air is discharged at the nozzle between inner flaps 8 and outer flaps 9. The inner flaps 8 are arranged as an extension of the rear part of the body of the turbojet engine 5, that is to say the afterburner duct 6. The outer flaps 9 are arranged as an extension of the housing 4 in which the turbojet engine 5 is located, that is to say as an extension of the aircraft's fuselage. The inner flaps 8 are therefore in contact with the high temperature propulsive gases and, for this reason, are called "hot flaps". The outer flaps 9 are located on the outside of the cooling air stream emerging from the duct 7, and for this reason are called "cold flaps". The cooling air therefore emerges from the duct 7 between the inner flaps 8 and the outer flaps 9, providing thermal protection for the surrounding rear part of the aircraft.
A seal 10 is provided between the rear end of the fuselage and the outer flaps 9, which are mounted so as to pivot about axes fixed in relation to the turbojet engine 5. This seal is necessary, inter alia, to protect the aircraft structure from possible back surges of combustion gases upstream of the nozzle. The seal assembly used for this purpose must be flexible because radial and axial displacements on the turbojet engine 5 occur relative to the aircraft's fuselage. In fact, there can arise axial displacements of the order of 20 mm, due to differences in temperature and in the materials used. Radial displacements may be of the order of 10 mm, and occur mainly during maneuvers such as tight turns or arrested landings. In addition, it may be noted that the outer flaps can have a travel of 10.degree..
In FIG. 1, the seal 10 between the fuselage and the outer flaps is shown in heavy lines, and represents equally the known sealing assemblies of the prior art and that which is the subject of the present invention.
The purpose of FIG. 1 is only to indicate the position of this sealing assembly in a military aircraft.
2. Description of the Prior Art
One known form of such a sealing assembly is shown in FIG. 2 of the drawings.
This known seal comprises an annular, flexible, metal front part 11 having its front end 12 forming an inwardly curved flexible portion in contact with a slide shoe 13 fixed to the fuselage of the aircraft 2. This front part 11 is preferably formed by flexible strips of titanium riveted to support legs 19 which are themselves each fixed to the periphery of the afterburner duct 6 of the turbojet engine 5. The end 12 of this front part 11 is preferably made of teflon to facilitate axial sliding movement of the seal.
The seal further comprises an intermediate part 21 which is of metal and is located to the rear of the support legs 19. This intermediate part 21 carries shoes 16, preferably of teflon, which contact the rear end 15 of the fuselage of the aircraft 2, thus permitting the entirety of the turbojet engine 5 to slide axially relative to the inner wall of the fuselage.
The seal is completed by a metal rear part 17 which is divided by longitudinal slits, each divided portion being in contact with the base of an outer flap 9.
The fact that this seal assembly consists of several split parts, which themselves are composed of several flexible parts, enables radial displacements due to distortion of the different parts of the seal to be accommodated.
The number of parts constituting such a seal is in the neighborhood of 250 parts, including rings, strips, bolts, rivets, of which a large number are made of titanium, and slider pads made of teflon. The assembly and mounting of such a seal is therefore complicated and its production cost is high. In addition, the large number of parts makes its weight far from negligible.